The objective of this work is to examine the flow structure and heat transfer distribution of a squealer tip rotor blade with purge flow cooling and provide a comparison with a basic flat tip rotor blade without purge flow cooling, under transonic conditions and high inlet free stream turbulence intensity. The blade design was provided by Solar Turbines Inc., and consists of a double squealer around the pressure and suction sides, two purge flow blowing holes located downstream of the leading edge and mid-chord, four ribs in the mid-chord region and a trailing edge bleeder exiting on the pressure side. Blade cavity depth is 2.29 mm (0.09 in.) and the total blade turning angle is 107.5°. Tests were performed in a blow-down facility at a turbulence intensity of 12%, in a seven bladed 2-D linear cascade at transonic conditions. Experiments were conducted at isentropic exit Mach numbers of 0.85 and 1.05, corresponding to Reynolds numbers based on axial chord of 9.75x10^5 and 1.15x10^6, respectively, and tip clearance gaps of 1% and 2% of the scaled engine blade span. A blowing ratio of 1.0 was used in the squealer tip experiments. Detailed heat transfer coefficient and film cooling effectiveness distributions were obtained using an infrared thermography technique, while oil flow visualization was used to investigate the flow patterns in the blade tip region. With the addition of a squealer tip, leakage flow was found to decrease, as compared to a flat tip blade. With increasing tip clearance gap, the heat transfer coefficients within the cavity and along the squealer rim were found to decrease and increase, respectively. Film cooling effectiveness decreased with increasing tip clearance gap and was mainly observed within the squealer cavity. The maximum heat transfer coefficient was observed on the leading edge, however, comparatively large values were observed on the mid-chord ribs. The presence of the ribs, greatly affected the flow structure and heat transfer distributions within the cavity and downstream towards the trailing edge. / Master of Science

This paper presents details of an experimental and computational investigation on the effect of misalignment between the combustor exit and nozzle guide vane endwall on the heat transfer distribution across an axisymmetric converging endwall. The axisymmetric converging endwall investigated was representative of that found on the shroud side of a first stage turbine nozzle section. The experiment was conducted at a nominal exit M of 0.85 and exit Re 1.5 x 10⁶ with an inlet turbulence intensity of 16%. The experiment was conducted in a blowdown transonic linear cascade wind tunnel. Two different inlet configurations were investigated. The first configuration, Case I, was representative of a combustor exit aligned to the nozzle platform, with a gap located at the interface of the tow components. The second configuration, Case II, the endwall platform was offset in the span-wise direction to create a backward facing step at the inlet. This step is representative of a misalignment between the combustor exit and the NGV platform. An infrared camera was used to capture the temperature history on the endwall, from which the endwall heat transfer distribution was determined. A numerical study was also conducted by solving RANS equations using ANSYS Fluent v.15. The numerical results provided insight into the passage flow field which explained the observed heat transfer characteristics. Case I showed the typical characteristics of transonic vane cascade flow, such as the separation line, saddle point, and horseshoe vortices. The presence of a gap at the combustor-nozzle interface facilitated the formation of a separated flow which propagated through the passage. This flow feature caused the passage vortex reattach to the SS vane at 0.44 x/C. The addition of the platform misalignment in Case II caused the flow reattachment region to occur near the vane LE plane. The separated flow which formed at the inlet step, merged with the recirculation region on the endwall platform, forming two counter-rotating auxiliary vortices. These vortices significantly delayed migration of the passage vortex, causing it to reattach on the SS vane at 0.85 x/C. These two flow features also had a significant effect on the endwall heat transfer characteristics. The heat transfer levels on the endwall platform, from -0.50 to +0.50 Cx relative to the vane LE, had an average increase of ~40%. However, downstream of the vane mid-passage, the heat transfer levels showed no appreciable heat transfer augmentation due to flow acceleration through the passage throat. / Master of Science

Modern gas turbine designs' ever-increasing turbine inlet temperature raises challenges for the nozzle guide vane cooling. Two typical endwall cooling schemes, jump cooling and louver cooling, result in different interactions between the injected coolant and the mainstream, leading to different cooling effects. This study investigates these two cooling schemes on the endwall cooling experimentally and numerically. Wind tunnel tests and the CFD simulations are carried out with engine-representative conditions of an exit Mach number of 0.85, an exit Reynolds number of 1.5×10^6, and an inlet Turbulence intensity of 16%. The jump cooling scheme experiments investigate two blowing ratios, 2.5 and 3.5, two density ratios, 1.2 and 1.95, and three endwall profiles with different NGV-turbine alignments. Four coolant mass flow ratios from 1.0% to 4.0% are tested for the louver cooling. The results show that the cavity vortex, the horseshoe vortex, and the passage vortex are the main factors that prevent the upstream coolant from reaching the NGV passage. The jump cooling scheme generally provides high momentum to the cooling jets. As a result, the coolant at the design case density ratio of 1.95 and blowing ratio of 2.5 is sufficiently energized to penetrate the horseshoe vortex. It then forms a relatively uniform coolant film near the NGV passage inlet, leading to a minimum adiabatic cooling effectiveness of 0.4 throughout the passage. Reducing the coolant density or increasing the blowing ratio leads to higher coolant momentum, so the coolant jets can further suppress the horseshoe vortex. However, high momentum may cause coolant lift-off, mitigating the coolant reattachment. Therefore, the density ratio needs to be carefully balanced with the blowing ratio to optimize the cooling effect. This balance is also affected by the combustor-NGV misalignment, as a higher step height requires higher coolant momentum to overcome the step-induced vortices. On the contrary, the louver cooling scheme provides less momentum to the coolant. The results showed that only by exceeding a coolant mass flow rate of 1~2% can the coolant form a uniform film which provides good coverage upstream of the NGV passage inlet. As for the cooling of the NGV passage, the mass flow rate ratio of the range investigated is not sufficient for desirable cooling performance. The pressure side endwall proves most difficult for the coolant to reach. In addition, the fishmouth cavity at the combustor-NGV passage causes a three-dimensional cavity vortex that transports the coolant in the pitch-wise direction. Moreover, the coolant transport pattern is dependent on the coolant blow rate. Overall, the more-energized coolant film generated by the jump cooling tends to survive longer, but it is also more prone to lift-off. At the same time, the less-energized coolant film caused by the louver cooling is more susceptible to vortices and the discontinuity of the endwall geometry. However, it develops faster, especially in the lateral direction. The two schemes could be applied simultaneously for an ideal cooling system. The jump cooling can provide enough momentum for the coolant to persist in the NGV passage. Meanwhile, the louver cooling covers the upstream region before the jump cooling coolant reattaches to the endwall. / Doctor of Philosophy / Gas turbines, sometimes called combustion turbines, are widely used to generate power or propulsion for various applications. The three main components of a gas turbine are compressor, combustor, and turbine. Modern gas turbines run at a high turbine inlet temperature that exceeds the current metal limits to increase efficiency. However, this brings significant challenges to the cooling of the first stage of the turbine, the nozzle guide vane. In this research, two commonly used endwall cooling methods, jump cooling and louver cooling, are investigated under engine-representative conditions experimentally and numerically. In addition, flow physics is demonstrated to explain the endwall cooling performance, mainly the upstream boundary layer caused by the interaction between the mainstream and the coolant flow. The results show that the cavity vortex, the horseshoe vortex, and the passage vortex are the main factors that prevent the upstream coolant from reaching the NGV passage. The jump cooling scheme provides high momentum to the cooling jets. As a result, the coolant in the design case is sufficiently energized to penetrate the horseshoe vortex, providing a desirable cooling effect in the NGV passage. Reducing the density ratio or increasing the blowing ratio can help the coolant jets further suppress the horseshoe vortex but also causes more lift-off, which adversely affects the cooling performance. On the contrary, the louver cooling scheme provides less momentum to the coolant, forming a less energized coolant film. The lack of coolant causes the louver coolant film to provide good coverage immediately downstream of the louver scheme exit. However, due to unfavorable interaction with vortices and endwall discontinuity, the cooling effect decays quickly downstream. Overall, the more-energized coolant film generated by the jump cooling tends to survive longer, but it is also more prone to lift-off. At the same time, the less-energized coolant film caused by the louver cooling is more susceptible to vortices and the discontinuity of the endwall geometry. However, it develops faster, especially in the lateral direction. The two schemes could be applied simultaneously for an ideal cooling system to work mutually beneficially.

The study presented herein describes and analyzes a detailed experimental investigation of the effects of density ratio on endwall thermal performance at varying blowing rates for a typical nozzle guide vane platform purge jet cooling scheme. An axisymmetric converging endwall with an upstream doublet staggered cylindrical hole purge jet cooling scheme was employed. Nominal exit flow conditions were engine representative and as follows: {rm Ma}_{Exit} = 0.85, {rm Re}_{Exit,C_{ax}} = 1.5 times {10}^6, and large-scale freestream Tu = 16%. Two blowing ratios were investigated corresponding to the upper and lower engine extrema. Each blowing ratio was investigated amid two density ratios; one representing typical experimental neglect of density ratio, at DR = 1.2, and another engine representative density ratio achieved by mixing foreign gases, DR = 1.95. All tests were conducted on a linear cascade in the Virginia Tech Transonic Blowdown Wind Tunnel using IR thermography and transient data reduction techniques. Oil paint flow visualization techniques were used to gather quantitative information regarding the alteration of endwall flow physics due two different blowing rates of high-density coolant. High resolution endwall adiabatic film cooling effectiveness, Nusselt number, and Net Heat Flux Reduction contour plots were used to analyze the thermal effects. The effect of density is dependent on the coolant blowing rate and varies greatly from the high to low blowing condition. At the low blowing condition better near-hole film cooling performance and heat transfer reduction is facilitated with increasing density. However, high density coolant at low blowing rates isn't adequately equipped to penetrate and suppress secondary flows, leaving the SS and PS largely exposed to high velocity and temperature mainstream gases. Conversely, it is observed that density ratio only marginally affects the high blowing condition, as the momentum effects become increasingly dominant. Overall it is concluded density ratio has a first order impact on the secondary flow alterations and subsequent heat transfer distributions that occur as a result of coolant injection and should be accounted for in purge jet cooling scheme design and analysis. Additionally, the effect of increasing high density coolant blowing rate was analyzed. Oil paint flow visualization indicated that significant secondary flow suppression occurs as a result of increasing the blowing rate of high-density coolant. Endwall adiabatic film cooling effectiveness, Nusselt number, and NHFR comparisons confirm this. Low blowing rate coolant has a more favorable thermal impact in the upstream region of the passage, especially near injection. The low momentum of the coolant is eventually dominated and entrained by secondary flows, providing less effectiveness near PS, near SS, and into the throat of the passage. The high momentum present for the high blowing rate, high-density coolant suppresses these secondary flows and provides enhanced cooling in the throat and in high secondary flow regions. However, the increased turbulence impartation due to lift off has an adverse effect on the heat load in the upstream region of the passage. It is concluded that only marginal gains near the throat of the passage are observed with an increase in high density coolant blowing rate, but severe thermal penalty is observed near the passage onset. / Master of Science / Gas turbine technology is used frequently in the burning of natural gas for power production. Increases in engine efficiency are observed with increasing firing temperatures, however this leads to the potential of overheating in the stages following. To prevent failure or melting of components, cooler air is extracted from the upstream compressor section and used to cool these components through various highly complex cooling schemes. The design and operational adequacy of these schemes is highly subject to the mainstream and coolant flow conditions, which are hard to represent in a laboratory setting. This experimental study explores the effects of various coolant conditions, and their respective response, for a purge jet cooling scheme commonly found in engine. This scheme utilizes two rows of staggered cylindrical holes to inject air into the mainstream from platform, upstream of the nozzle guide vane. It is the hope that this air forms a protective layer, effectively shielding the platform from the hostile mainstream conditions. Currently, little research has been done to quantify these effects of purge flow cooling scheme while mimicking engine geometry, mainstream and coolant conditions. For this study, an endwall geometry like that found in engine with a purge jet cooling scheme is studied. Commonly, an upstream gap is formed between the combustor lining and first stage vane platform, which is accounted for in this testing. Mainstream and coolant flow conditions can have large impacts on the results gathered, so both were matched to engine conditions. Varying of coolant density and injection rate is studied and quantitative results are gathered. Results indicate coolant fluid density plays a large role in purge jet cooling, and with neglection of this, potential thermal failure points could be overlooked This is exacerbated with less coolant injection. Interestingly, increasing the amount of coolant injected decreases performance across much of the passage, with only marginal gains in regions of complex flow. These results help to better explain the impacts of experimental neglect of coolant density, and aid in the understanding of purge jet coolant injection.

The flow and heat transfer within compressor rotor cavities of aero-engines is a conjugate problem. Depending on the operating conditions buoyancy forces, caused by radial temperature difference between the cold throughflow and the hotter shroud, can influence the amount of entrained air significantly. By this, the heat transfer depends on the radial temperature gradient of the cavity walls and in reverse the disk temperatures are dependent on the heat transfer. In this thesis, disk Nusselt numbers are calculated in reference to the air inlet temperature and in comparison to a modeled local air temperature inside the cavity. The local disk heat flux is determined from measured steady-state surface temperatures by solving the inverse heat transfer problem in an iterative procedure. The conduction equation is solved on a 2D mesh, using a validated finite element approach and the heat flux confidence intervals are calculated with a stratified Monte Carlo approach. An estimate for the amount of air entering into the cavity is calculated by a simplified heat balance. In addition to the thermal characterization of the cavity, the mass exchange of the air in the cavity with the axial flow in the annular gap and the swirl distribution of the air in the cavity are also investigated.:1 Einleitung 2 Grundlagen und Literaturübersicht 2.1 Modellsystem der rotierenden Kavitäten mit axialer Durchströmung 2.2 Ergebnisgrößen 2.3 Strömung in rotierenden Kavitäten 2.4 Wärmeübertragung in rotierenden Kavitäten 2.5 Fluidtemperatur in rotierenden Kavitäten 3 Experimenteller Aufbau 4 Messtechnik 4.1 Oberflächen- und Materialtemperaturen 4.2 Lufttemperaturen 4.3 Statischer Druck 4.4 Dreiloch-Drucksonden 5 Datenauswertung 5.1 Kernrotationsverhältnis 5.2 Wärmestromdichte und Nusseltzahl 5.2.1 Finite-Elemente Modell 5.2.2 inverses Wärmeleitungsproblem 5.2.3 Anpassungsmethode 5.2.4 Testfälle zur Validierung 5.2.5 Validierung Testfall 1 und 3 - ideale Kavitätenscheibe 5.2.6 Validierung Testfall 2 - Reproduzierbarkeit 5.2.7 Validierung Testfall 4 - lokales Ereignis 5.2.8 Bestimmung der Wärmestromdichte-Unsicherheit 5.2.9 Anwendung der Anpassungsmethode auf experimentelle Daten 5.2.10 Wahl der Randbedingungsfunktion 5.2.11 Wärmeübergangskoeffizient und Nusselt-Zahl 5.2.12 Zusammenfassung 5.3 Austauschmassenstrom 6 Experimentelle Ergebnisse 6.1 Dichteverteilung in der Kavität 6.2 Massenaustausch Kavität 6.3 Wärmeübertragung in der Kavität 6.3.1 Fallbeispiel 6.3.2 Einfluss der Drehfrequenz 6.3.3 Einfluss des Massenstromes 6.3.4 Einfluss des Auftriebsparameters 6.4 Wärmeübertragung im Ringspalt 6.5 Drall im Ringspalt und der Kavität 7 Zusammenfassung und Ausblick / Die Strömung und Wärmeübertragung in den Verdichterkavitäten von Flugtriebwerken ist ein konjugiertes Problem. Durch die radialen Temperaturunterschiede in der Kavität wird die Menge der in die Kavität strömenden Luft stark beeinflusst. Somit ist die Wärmeübertragung abhängig von den radialen Temperaturgradienten der Scheibenwände und umgekehrt ist die Scheibentemperatur abhängig von der Wärmeübertragung. Die Nusselt-Zahl in diesem System wurde aufgrund der schwierigen Zugänglichkeit in der Historie auf die eine Referenztemperatur vor der Kavität bezogen. Dies ist insofern problematisch, da hierdurch die thermischen Verhältnisse unterschätzt werden können. In dieser Arbeit wird ein neuer Ansatz zu Berechnung der Nusselt-Zahl mithilfe einer modellierten lokalen Lufttemperatur innerhalb der Kavität verwendet. Die lokale Wärmestromdichte auf der Scheibenoberfläche wird mithilfe eines validierten zweidimensionalen rotationssymmetrischen Finite-Element Modells auf der Grundlage von gemessenen Oberflächentemperaturen berechnet. Dies stellt ein inverses Wärmeleitungsproblem dar, welches mithilfe einer Anpassungsmethode gelöst wurde. Die Auswirkung der Messunsicherheit der Temperaturmessung auf die berechnete Wärmestromdichte wird durch eine geschichtete Monte-Carlo-Simulation, nach dem Ansatz der LHC-Methode, untersucht. Neben der thermischen Charakterisierung der Kavität wird zudem der Massenaustausch der Luft in der Kavität mit der axialen Durchströmung im Ringspalt sowie die Drallverteilung der Luft in der Kavität untersucht.:1 Einleitung 2 Grundlagen und Literaturübersicht 2.1 Modellsystem der rotierenden Kavitäten mit axialer Durchströmung 2.2 Ergebnisgrößen 2.3 Strömung in rotierenden Kavitäten 2.4 Wärmeübertragung in rotierenden Kavitäten 2.5 Fluidtemperatur in rotierenden Kavitäten 3 Experimenteller Aufbau 4 Messtechnik 4.1 Oberflächen- und Materialtemperaturen 4.2 Lufttemperaturen 4.3 Statischer Druck 4.4 Dreiloch-Drucksonden 5 Datenauswertung 5.1 Kernrotationsverhältnis 5.2 Wärmestromdichte und Nusseltzahl 5.2.1 Finite-Elemente Modell 5.2.2 inverses Wärmeleitungsproblem 5.2.3 Anpassungsmethode 5.2.4 Testfälle zur Validierung 5.2.5 Validierung Testfall 1 und 3 - ideale Kavitätenscheibe 5.2.6 Validierung Testfall 2 - Reproduzierbarkeit 5.2.7 Validierung Testfall 4 - lokales Ereignis 5.2.8 Bestimmung der Wärmestromdichte-Unsicherheit 5.2.9 Anwendung der Anpassungsmethode auf experimentelle Daten 5.2.10 Wahl der Randbedingungsfunktion 5.2.11 Wärmeübergangskoeffizient und Nusselt-Zahl 5.2.12 Zusammenfassung 5.3 Austauschmassenstrom 6 Experimentelle Ergebnisse 6.1 Dichteverteilung in der Kavität 6.2 Massenaustausch Kavität 6.3 Wärmeübertragung in der Kavität 6.3.1 Fallbeispiel 6.3.2 Einfluss der Drehfrequenz 6.3.3 Einfluss des Massenstromes 6.3.4 Einfluss des Auftriebsparameters 6.4 Wärmeübertragung im Ringspalt 6.5 Drall im Ringspalt und der Kavität 7 Zusammenfassung und Ausblick